A typical aircraft gas turbine engine includes, in serial flow relationship, an engine inlet, a fan, a core engine, including a compressor, combustor, and high and low pressure turbines, defining a core flowpath for core airflow, and an exhaust nozzle for accelerating the core airflow to obtain maximum thrust. In a turbofan type gas turbine engine a portion of fan air bypasses the core and rejoins the core airflow aft of the turbines and forward of the exhaust nozzle. In a turbojet engine there is no fan or bypass. An exhaust nozzle for an engine designed for supersonic operation typically includes a convergent section for accelerating the core airflow to choked flow (Mach 1) at the nozzle throat, and a divergent section for accelerating the airflow supersonically. An ejector can be provided in the divergent section whereby the higher velocity/higher pressure core gas stream is used to withdraw lower velocity/lower pressure gas stream, conventionally from a source of ram or ambient air. Such ejecting exhaust nozzles have been developed for both cooling and sound suppression purposes. U.S. Pat. No. 3,463,402, Jet Sound Suppressing Means, to Charles E. Langston, Jr., teaches use of an ejecting exhaust nozzle to generate turbulence between two streams of gas, thus suppressing the noise normally generated from shear stresses at the interface of the higher velocity core gas stream with the lower velocity gas stream. U.S. Pat. No. 3,409,228, Ejector Nozzle, to Hans P. Mehr, teaches an ejector nozzle for cooling a gas turbine engine and exhaust nozzle, whereby an ejector is in flow communication with an engine inlet and draws the ram airflow through a bypass duct to provide cooling to the engine and exhaust nozzle, additionally providing a ram air inlet proximate the exhaust nozzle for low Mach number operation wherein the pressure drop through the bypass duct inhibits the desired ejector flow.
It has been found that proper ejector operation for withdrawing the lower pressure air through an ejector into the higher velocity/higher pressure gas stream is dependent on, for a given ejector configuration, the pressure ratio defined by the pressure of the ejected air supply divided by the core airflow pressure. FIG. 1 is a representative flight map of Mach number versus altitude for an aircraft with an augmented gas turbine engine. Contour lines are provided representing the pressure ratio defined by the available supply ram air pressure, P.sub.RAM, divided by the core airflow total pressure at the nozzle throat, P.sub.8, for the representative engine at maximum power. The nozzle throat is conventionally referred to as station 8, thus the corresponding pressure P.sub.8 is a conventional parameter used by those skilled in the art of exhaust nozzle design and performance. The nozzle throat is also a conventional location for an ejector, as the core airflow static pressure drops and velocity increases beginning at the throat when the flow is choked there. For a given ejector nozzle configuration, including size and location between the nozzle throat and nozzle exit, to operate from ram pressure only, the pressure ratio P.sub.RAM /P.sub.8 must be above a critical ejector pressure ratio, determined by conventional methods, for that particular ejector configuration. If, for example, the critical ejector pressure ratio is determined to be 0.22, the ejector would not operate unless the aircraft was operating to the fight of contour line 0.22, at a relatively high Mach number for a given altitude. If the pressure ratio is below the critical ejector pressure ratio for the configuration, flow through the ejector will be inhibited, and if too low, the ejector will be subject to flow in the reverse direction, i.e., from the core flow to the ejector air source. At low Mach numbers, the ram air pressure is low because the velocity component of the total ram pressure is negligible, and could be below the critical value. If this occurs, the ejector will not operate to pump air into the core stream for either cooling or sound suppression. In modern military gas turbine engines which have a high engine pressure ratio (P.sub.8 /P.sub.RAM), it is difficult to provide cooling air to the divergent exhaust nozzle, especially at low Mach numbers. In engines including a thrust augmentor, it is particularly critical to provide cooling airflow to the exhaust nozzle whenever the augmentor is on, including at low Mach numbers. Higher pressure fan air is therefore often used for nozzle cooling, as cooling from a ram air ejector alone could be inadequate. However, use of such fan air is quite expensive in terms of engine performance and efficiency.